Fiber-oriented repair geometries for composite materials

Niedernhuber, M.; Holtmannspötter, Jens; Ehrlich, Ingo

doi:10.1016/j.compositesb.2016.03.027

Cited by 24 publications

(17 citation statements)

References 16 publications

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“…This results in reduced material removal and a repair area up to 50% smaller than a conventional step design, with theoretically equivalent strength. Initial mechanical test results, for representative joints, have shown that the tensile strength of fibre-oriented step joints can match that of traditional step joints that are 40% longer, both with and without over-plies [14]. However, even with a very shallow taper (50:1 or 1.145°), these joints were only capable of restoring 59% and 54% of the pristine tensile strength of the parent laminate with and without over-plies, respectively.…”

Section: Repair Design Optimisationmentioning

confidence: 99%

“…Recent work by Niedernhuber et al [14] proposed that composite step repairs may be optimised by employing a fibre-oriented step approach, where step lengths for each ply are shortened based on the orientation of fibres in each ply relative to the bondline direction, as shown in Figure 1. This relationship is summarised in Equation 1, which defines the appropriate step length for any fibre-oriented ply, , based on the fibre direction in that ply, , the ply thickness, , and the ideal minimum taper angle, , that would be used for a conventional step approach.…”

Section: Fibre-oriented Step Designmentioning

confidence: 99%

“…Next the two arc/irregular segments, 1 and 2 , can be defined based on the polar profile of the previous ply, −1 and −1 , and the fibre direction of the current ply, , according to Equations (14) and (15). This is reflected by step 4 of Figure 4.…”

Section: = Tanmentioning

confidence: 99%

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Modelling the size and strength benefits of optimised step/scarf joints and repairs in composite structures

Pierce

Falzon

2019

Composites Part B: Engineering

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Adhesive bonding offers a better load transfer between adherends, for the assembly and repair of composite structures, compared with mechanical fastening methods. One drawback of adhesive bonded repairs is the considerable amount of material which must be removed, around the damaged region, to ensure adequate load transfer. Optimised geometries that account for the highly-orthotropic properties of individual composite plies are investigated, including a novel 'fibre-oriented' scarf approach that is inspired by an existing fibre-oriented step design. These methods aim to reduce the length of joint and repair bond-lines by at least 36%, compared with conventional step and scarf geometries of a similar strength. Size-reduction benefits are predicted using a MATLAB tool that is applicable for any composite laminate, and parametric analysis used to assess the effect of ply thickness, the number of plies, stacking sequence and taper angle. Cohesive Zone Models of joints and repairs with conventional and fibre-oriented designs are used to predict and compare the ultimate strength of each configuration. The existing fibre-oriented step design appears to show no benefit over a conventional step design. However, the novel fibre-oriented scarf approach results in a 33-40% bond-line length reduction compared to a conventional scarf design with similar strength. Analysis further indicates a 17-22% increase in ultimate strength for joints and repairs with the same bond-line length that employ the optimised fibre-oriented scarf design.

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Section: Repair Design Optimisationmentioning

confidence: 99%

Section: Fibre-oriented Step Designmentioning

confidence: 99%

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Modelling the size and strength benefits of optimised step/scarf joints and repairs in composite structures

Pierce

Falzon

2019

Composites Part B: Engineering

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“…CFRP materials have for many years been used for the primary structures of light aeroplanes, gliders and military aircraft. Only in recent years with the advent of civil transport aircraft such as the Boeing 787, the Airbus A350 XWB and A220 the application of CFRP materials in civilian aircraft has evolved from secondary structures such as fairings to primary and load bearing structures such as the wing box and fuselage (1,2). A significant difference between light aeroplanes and large airliners is the change from a monocoque to a stressed-skin construction, enabling high levels of structural loading and structural efficiency (3).…”

Section: Introductionmentioning

confidence: 99%

Understanding the influence of laminate stacking sequence on strain/stress concentrations in thin laminates at repair holes with large scarf angles

Damghani

Bakunowicz

Murphy³

2019

Journal of Composite Materials

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Scarf repair is widely used in the restoration of structural performance of damaged aircraft secondary structures. Such repairs result in reduced thickness sections which are significantly larger than those associated with typical fastener holes. Significant literature exists on the distribution of strain/stress concentration in fastener hole geometries, both straight sided and countersunk, but is lacking for the geometries associated with shallow scarf angles and thin laminates. Hence, herein three-dimensional finite element models are developed to understand the influence of stacking sequence and scarf angle on strain/stress concentrations. The results demonstrate and quantify for the first time that strain concentrations are not only dependant on the laminate membrane stiffness but also on laminate bending stiffness, due to the anisotropy created as a result of scarfing angle, hole geometry and laminate thickness. Scarfing is demonstrated, for typical repair geometry associated with foreign object damage (hole diameter 20 mm, scarf angles 3° to 7°), to elevate strains by up to 2.5 times when compared to equivalent diameter straight-sided holes in laminates of thickness ≈1 mm.

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“…Soft patch technology was chosen for the repair. As recommended [3], restoration of structural continuity in terms of reinforcement directions was attempted wherever possible. To achieve this, the lay-up of the fabric constituting the patch was chosen as shown in Fig.…”

Section: Introductionmentioning

confidence: 99%

Field repair of a severely damaged FG/epoxy fuselage

Czarnocki¹,

Zagrajek²,

Switkiewicz³

2019

MATEC Web Conf.

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The paper presents development of a repair procedure for a severely damaged primary laminate structure of a light aircraft, certified in accordance with CS22. The procedure was developed with the assumption that it would be a field repair carried out by people of average manual skills, who have not been trained in repairs of composite structures. To develop the procedure, a typical FG/epoxy laminate glider fuselage was used. The primary objective of the repair was to restore the original fuselage stiffness, while maintaining possibly low stress concentrations resulting from the repairs. To facilitate development of the repair, an FE model was developed. It was parametrized to be useful for design of similar repairs of similar structures. The experimental work done with the use of a real structure allowed the authors to verify numerical simulations, which were found to be correct.

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Fiber-oriented repair geometries for composite materials

Cited by 24 publications

References 16 publications

Modelling the size and strength benefits of optimised step/scarf joints and repairs in composite structures

Modelling the size and strength benefits of optimised step/scarf joints and repairs in composite structures

Understanding the influence of laminate stacking sequence on strain/stress concentrations in thin laminates at repair holes with large scarf angles

Field repair of a severely damaged FG/epoxy fuselage

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